Rocket engine liquid fueled thrust chambers of the larger thrust type, typically of many hundreds, many thousands or even millions of pounds of thrust, employ regeneratively cooled thrust chambers where pressurized (pumped) propellant is first passed through thrust engine tubing or channels forming the shell or cooling jacket of the chamber, before being injected into the combustion chamber. The cool fuel or oxidizer in a bipropellant system, for example, liquid oxygen at -180.degree. C. or other oxidizer, thus keeps the combustion chamber at a sufficiently low temperature to preserve the structural integrity of the thrust chamber. In the case of smaller thrust thrusters used for intermittent thrust control of a space vehicle or satellite, thrust chambers have employed film cooling. Film cooling employs a protective coating of propellant which is sprayed along the inner surface of the thrust chamber. Evaporation of the film cools the chamber wall. Although film cooling is efficient, it is to be avoided since it lowers the overall specific thrust by using propellant for a purpose other than producing thrust. Regeneratively cooled engines are considered more efficient since coolant is not wasted but, in fact, augments the initial energy at injection by its increased heat content.
Conventional thrusters currently in use have a minimal upper temperature limit of about 2400.degree. F. (1315.degree. C.) and a limited life span of about ten hours. These conventional thrusters, using a hydrazine propellant for example, and a thrust chamber constructed of niobium alloys, necessarily will use about 40% of the fuel for film cooling in order to keep the thrust chamber walls below this temperature. Since the propellant is the major mass item for satellites being put in space, a considerable incentive exists to decrease or obviate the need for film cooling and hence the amount of on-board fuel.
U.S. Pat. No. 3,354,652 discusses the difficulty of regeneratively cooling small liquid propellant engines resulting inter alia in boiling or decompositon of the coolant within the coolant jacket. While it has been suggested to apply high temperature insulation, e.g., metal oxides, to the combustion side of the chamber to reduce the coolant bulk temperature during steady state firing, this can result, upon engine shut down, in additional stored heat in the insulation causing localized heating and decomposition of remaining stagnant propellant. The patent somewhat solves the problem by suggesting a tantalum alloy liner coupled with a stagnant gas or vacuum enclosed space and helical two-way flow coolant channels.
U.S. Pat. No. 3,780,533 discloses the use in regeneratively cooled chambers utilizing cooling channels, of a composite wall including a deposit of electroformed nickel, or a sheet of nickel or of refractory alloys, such as copper-silver or molybdenum-rhenium alloys, brazed to lands in a middle wall component U.S. Pat. No. 3,315,471 shows with respect to thrusters utilizing radioisotope fuel, structural elements of the thruster, namely spaced shells, preferably constructed of tungsten. U.S. Pat. No. 3,723,742 shows the use of noble metals and refractory metals surrounding a radioisotope fuel casing.
U.S. Pat. No. 4,917,968 decribes a thrust chamber structure where a ductile layer of a platinum group metal including iridium is deposited by chemical vapor deposition on a mandrel and a layer of refractory metal deposited thereover also by chemical vapor deposition, with a solid solution of the two metals present between and metallurgically bonded to the two metal layers.
U.S Pat. No. 5,613,299 describes a thrust chamber structure where a layer of a platinum group metal, including iridium, is bonded to the interior of a refractory alloy thrust chamber by pressurizing the exterior of the chamber forcing the chamber to collapse onto the liner, itself supported on a solid mandrel. This present invention differs from the above by having a hollow mandrel so that the interior of the chamber is pressurized and the liner expanded outwards.